The present invention relates generally to gas turbine engines, and, more specifically, to deicing therein.
Turbofan aircraft gas turbine engines are configured for powering an aircraft in flight from takeoff, cruise at altitude, descent, and landing in various weather conditions. Since the temperature of operation varies substantially and includes below-freezing temperatures, the engine is subject to inlet icing conditions.
Humid air and rain may enter an engine in freezing conditions and may deposit as layers of ice on internal components of the engine. For example, in icing operating conditions ice may accumulate on fan blades, inlet guide vanes (IGVs) of the booster compressor, annular splitter between the booster compressor and the fan bypass duct, and along outlet guide vanes (OGVs) of the fan bypass duct. Since the splitter diverts flow into the booster compressor and the fan bypass duct, it is subject to a substantial amount of ice accumulation in specific icing conditions.
Since icing changes the aerodynamic profiles of the components being iced, the aerodynamic performance of the engine is adversely affected. And, liberated ice may be ingested through the booster compressor and additionally affect performance.
Ice accumulation is conventionally accommodated by configuring affected compressor airfoils with an increase in ruggedness to avoid or minimize damage from ice liberation. And, engine operability performance may be corrected by raising flight idle or ground idle speeds without violating corresponding thrust constraints.
However, in designing turbofan aircraft engines with even higher bypass ratios, the engine operability issues become more severe than previously encountered since more engine airflow will correspondingly increase the amount of ice accumulation which must be accommodated.
Furthermore, ever larger fan blades are being designed with state of the art composite materials and operate at slower rotational speeds. Slow fan speed permits more accumulation of ice in the specific icing conditions, which ice is shed at correspondingly higher rotational speed and increases ice damage potential of the downstream compressor components.
In one type of low bypass ratio military engine used in this country for many years, ice accumulation is reduced by providing hot compressor bleed air suitably channeled radially through fan front frame struts disposed in front of corresponding variable inlet guide vanes (VIGVs) in front of the first stage fan blades. The hot compressor air heats the front struts for deicing thereof when required in the flight envelope, which struts are otherwise unheated during the remainder of flight envelope operation.
However, strut heating requires substantially larger or thicker struts for channeling the hot bleed air therethrough, and significantly decreases the aerodynamic performance of the engine. Engine airfoil components such as the front frame struts are designed with specific aerodynamic profiles. Those profiles should be as small and thin as possible for maximizing aerodynamic performance, yet internal heating thereof requires the struts to be hollow and thicker than they otherwise would.
Accordingly, it is desired to provide a turbofan high bypass gas turbine engine with deicing capability for the booster compressor.
A booster compressor includes inlet guide vanes supported from a shroud. A shell surrounds the shroud and defines a manifold. A splitter nose includes a groove receiving a forward tang of the shroud with a clearance therebetween defining an outlet for the manifold. Hot air is channeled through the manifold and out the splitter nose for deicing thereof.